US3129561A - Rocket engine igniter - Google Patents

Rocket engine igniter Download PDF

Info

Publication number
US3129561A
US3129561A US54966A US5496660A US3129561A US 3129561 A US3129561 A US 3129561A US 54966 A US54966 A US 54966A US 5496660 A US5496660 A US 5496660A US 3129561 A US3129561 A US 3129561A
Authority
US
United States
Prior art keywords
matrix
engine
propellant
pellets
igniter
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
US54966A
Inventor
Joseph J Priapi
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Raytheon Technologies Corp
Original Assignee
United Aircraft Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Aircraft Corp filed Critical United Aircraft Corp
Priority to US54966A priority Critical patent/US3129561A/en
Application granted granted Critical
Publication of US3129561A publication Critical patent/US3129561A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42C—AMMUNITION FUZES; ARMING OR SAFETY MEANS THEREFOR
    • F42C19/00—Details of fuzes
    • F42C19/08—Primers; Detonators
    • F42C19/0807—Primers; Detonators characterised by the particular configuration of the transmission channels from the priming energy source to the charge to be ignited, e.g. multiple channels, nozzles, diaphragms or filters
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/95—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by starting or ignition means or arrangements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42C—AMMUNITION FUZES; ARMING OR SAFETY MEANS THEREFOR
    • F42C19/00—Details of fuzes
    • F42C19/08—Primers; Detonators
    • F42C19/0819—Primers or igniters for the initiation of rocket motors, i.e. pyrotechnical aspects thereof
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42C—AMMUNITION FUZES; ARMING OR SAFETY MEANS THEREFOR
    • F42C19/00—Details of fuzes
    • F42C19/08—Primers; Detonators
    • F42C19/12—Primers; Detonators electric

Definitions

  • This invention relates to an improved solid propellant gn tion system. It is designed to produce satisfactory 1gn1t1on for solid propellant rocket motors over a wide range of operating conditions, to increase ignition reliability and reproducibility and to reduce ignition peak pressures.
  • the pellet igniter consists of a pyrotechnic device wherein pellets of an explosive mixture are ignited which drive hot gas and particles into the propellant, igniting the propellant. It is sometimes difiicult to secure uniform ignition when using this system and high peak ignition pressures are encountered.
  • the rocket engine type of igniter essentially con sists of a pressure vessel filled with a propellant mixture comprising a small rocket engine which produces a supersonic flow of hot gas against the propellant grain causing the grain to ignite. Hot particles (which would be highly desirable) are absent in the gas stream.
  • an igniter for solid propellant engine systems which combines the advantageous features of the pellet type of igniter with those of the rocket engine type of igniter but without the drawbacks of either system.
  • the invention is accomplished by providing pellets of a finely divided metal together with an oxidizer with the pellets held in a matrix of a rocket propellant.
  • the matrix is formed into the shape of a standard internal burning charge and there is no nozzle or restriction at the end thereof, so that the flow of gas is subsonic.
  • FIGURE 1 is a partial sectional view of the head-end of a rocket engine showing an igniter in place embodying the present invention.
  • FIGURE 2 is an enlarged perspective View, partly in section, showing one form of pellet which can be employed in fabricating the igniter.
  • FIGURE 3 is a view, similar to FIGURE 2, showing another form of pellet.
  • the initiator assembly 11 is threadably mounted in the adapter.
  • the initiator assembly has the electric squibs therein, connected by means of wires 12 to a source of electricity, not shown.
  • the initiator assembly comprises a first chamber 13 which is filled with an igniter mixture 14, such as black powder or igniter pellets. This is retained on the perforated plate 15.
  • the igniter itself consists of a generally cylindrical chamber 17 which is threadably mounted in the adapter ring 9.
  • the chamber 17 is filled with an ignition mixture 19 which is described hereinafter.
  • the mixture 19 can be placed in the chamber 17 in any of the convention configurations but is preferably formed into a standard internal burning, tapered charge configuration such as that shown.
  • the ignition mixture contains a matrix 21 of a propellant mixture in which are encased pellets 23 of suitable size and shape as is hereinafter described.
  • the pellets themselves form no part of the present invention since such ignition pellets have been long known.
  • the pellets consists of a mixture of a metal and an oxidizer.
  • the metal is in the form of a relatively fine powder while the oxidizer is formed as a matrix around the metal particles.
  • the metal can be boron and the oxidizer potassium nitrate or the particles of metal might be aluminum powder while the 5 oxidizer is ammonium perchlorate.
  • About one-fourth of the pellet is a finely divided metal, the balance being an oxidizer together with a small amount of a binder.
  • the pellets can be cylindrical as is shown in FIGURE 2 where a single pellet grain 23 is shown.
  • the metal parti- 10 cles 25 are surrounded by the oxidizer matrix 27.
  • the pellets are in the form of small cylinders one-eighth inch in diameter and three-sixteenth inch long.
  • the pellets 29 are in the form of a small sphere one-eighth to one-fourth inch in diameter.
  • the metal particles 31 are surrounded by the oxidizing matrix material 33.
  • the ignition composition contains a matrix material 21 which can be any suitable propellant mixture such as the double base or composite propellant systems including those employing a fuel-binder such as polyurethane,
  • a conventional oxidizer such as ammonium perchlorate or ammonium nitrate.
  • the pellets are incorporated in the propellant mixture, preferably in a proportion of 50 to 75% by weight.
  • the squibs 10 To initiate the ignition it is only necessary to supply electricity through the wires 12 to the squibs 10.
  • the squibs 10 in turn ignite the charge 14 blowing hot combustion products through the perforated plate 15 which ignites the propellant charge 19. This in turn gnites the main propellant charge 5.
  • An igniter for a solid propellant rocket engine said engine having a combustion chamber confining a solid propellant, said propellant having a passage extending therethrough co-axially with the axis of the engine, said passage extending the length of the engine so as to define an internal burning charge, said igniter comprising:
  • a matrix of a solid rocket propellant comprising a mixture of a fuel binder and an oxidizer having dispersed therein solid pellets, said pellets comprising finely divided metal and an oxidizer therefor, all of said pellets being dispersed within and surrounded by said matrix, said matrix being confined by said cylindrical open-ended chamber, said matrix having a passage extending therethrough co-axially with the axis of the engine;
  • said igniter providing a subsonic flow of hot gas having hot metallic particles therein for the ignition of the solid propellant of the rocket engine.

Landscapes

  • Engineering & Computer Science (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • Air Bags (AREA)

Description

ROG/(7 CASING 1 PROPELLA/VT Cl-MRGZ' INVENTOR. s/QSEP/l (.[fiQ/AP/ ATTORNEYS April 21, 1964 J. J. PRlAPl ROCKET ENGINE IGNITER Filed Sept. 9, 1960 United States Patent Ofi ice 3,129,561 Patented Apr. 21, 1964 3,129,561 ROCKET ENGINE IGNITER Joseph J. Priapi, Menlo Park, Califi, assignor to United AircraftCorporation, a corporation of Delaware Filed Sept. 9, 1960, Ser. No. 54,966 4 Claims. (Cl. 60--35.6)
This invention relates to an improved solid propellant gn tion system. It is designed to produce satisfactory 1gn1t1on for solid propellant rocket motors over a wide range of operating conditions, to increase ignition reliability and reproducibility and to reduce ignition peak pressures.
Heretofore, the most common methods of igniting solid propellant rockets have been either by a pellet type ign1ter or by a rocket engine type igniter. The pellet igniter consists of a pyrotechnic device wherein pellets of an explosive mixture are ignited which drive hot gas and particles into the propellant, igniting the propellant. It is sometimes difiicult to secure uniform ignition when using this system and high peak ignition pressures are encountered. The rocket engine type of igniter essentially con sists of a pressure vessel filled with a propellant mixture comprising a small rocket engine which produces a supersonic flow of hot gas against the propellant grain causing the grain to ignite. Hot particles (which would be highly desirable) are absent in the gas stream.
In accordance with the present invention, an igniter for solid propellant engine systems is provided which combines the advantageous features of the pellet type of igniter with those of the rocket engine type of igniter but without the drawbacks of either system. Generally speaking, the invention is accomplished by providing pellets of a finely divided metal together with an oxidizer with the pellets held in a matrix of a rocket propellant. The matrix is formed into the shape of a standard internal burning charge and there is no nozzle or restriction at the end thereof, so that the flow of gas is subsonic.
In the drawings forming part of this application:
FIGURE 1 is a partial sectional view of the head-end of a rocket engine showing an igniter in place embodying the present invention.
FIGURE 2 is an enlarged perspective View, partly in section, showing one form of pellet which can be employed in fabricating the igniter.
FIGURE 3 is a view, similar to FIGURE 2, showing another form of pellet.
Referring now to the drawings by reference characters, there is shown a portion of a rocket engine casing 3 having a propellant grain 5 therein. The casing has a boss 7 which serves to retain the adapter 9. The initiator assembly 11 is threadably mounted in the adapter. The initiator assembly has the electric squibs therein, connected by means of wires 12 to a source of electricity, not shown. The initiator assembly comprises a first chamber 13 which is filled with an igniter mixture 14, such as black powder or igniter pellets. This is retained on the perforated plate 15. The igniter itself consists of a generally cylindrical chamber 17 which is threadably mounted in the adapter ring 9. The chamber 17 is filled with an ignition mixture 19 which is described hereinafter. The mixture 19 can be placed in the chamber 17 in any of the convention configurations but is preferably formed into a standard internal burning, tapered charge configuration such as that shown. Generally Speaking the ignition mixture contains a matrix 21 of a propellant mixture in which are encased pellets 23 of suitable size and shape as is hereinafter described.
The pellets themselves form no part of the present invention since such ignition pellets have been long known. Generally speaking the pellets consists of a mixture of a metal and an oxidizer. The metal is in the form of a relatively fine powder while the oxidizer is formed as a matrix around the metal particles. As an example, the metal can be boron and the oxidizer potassium nitrate or the particles of metal might be aluminum powder while the 5 oxidizer is ammonium perchlorate. About one-fourth of the pellet is a finely divided metal, the balance being an oxidizer together with a small amount of a binder. The pellets can be cylindrical as is shown in FIGURE 2 where a single pellet grain 23 is shown. Here the metal parti- 10 cles 25 are surrounded by the oxidizer matrix 27. In this particular embodiment, the pellets are in the form of small cylinders one-eighth inch in diameter and three-sixteenth inch long. In that form as shown in FIGURE 3, the pellets 29 are in the form of a small sphere one-eighth to one-fourth inch in diameter. Again the metal particles 31 are surrounded by the oxidizing matrix material 33.
The ignition composition contains a matrix material 21 which can be any suitable propellant mixture such as the double base or composite propellant systems including those employing a fuel-binder such as polyurethane,
polybutacliene, acrylic acid and the like together with a conventional oxidizer such as ammonium perchlorate or ammonium nitrate. The pellets are incorporated in the propellant mixture, preferably in a proportion of 50 to 75% by weight.
To initiate the ignition it is only necessary to supply electricity through the wires 12 to the squibs 10. The squibs 10 in turn ignite the charge 14 blowing hot combustion products through the perforated plate 15 which ignites the propellant charge 19. This in turn gnites the main propellant charge 5.
I claim:
1. An igniter for a solid propellant rocket engine, said engine having a combustion chamber confining a solid propellant, said propellant having a passage extending therethrough co-axially with the axis of the engine, said passage extending the length of the engine so as to define an internal burning charge, said igniter comprising:
(a) an open-ended cylindrical chamber positioned at one end of said passage, said open end of said chamber being directed down the passage;
(1)) a matrix of a solid rocket propellant comprising a mixture of a fuel binder and an oxidizer having dispersed therein solid pellets, said pellets comprising finely divided metal and an oxidizer therefor, all of said pellets being dispersed within and surrounded by said matrix, said matrix being confined by said cylindrical open-ended chamber, said matrix having a passage extending therethrough co-axially with the axis of the engine;
(0) ignition means for said matrix positioned at one end thereof; and
(d) said igniter providing a subsonic flow of hot gas having hot metallic particles therein for the ignition of the solid propellant of the rocket engine.
2. The igniter assembly of claim 1 wherein the pellets contain a mixture of boron and potassium nitrate.
3. The igniter of claim 1 wherein the pellets comprise a mixture of aluminum and ammonium perchlorate.
4. The igniter of claim 1 wherein the pellets comprise from 50 to 75 of the igniter mixture.
References Cited in the file of this patent UNITED STATES PATENTS

Claims (1)

1. AN IGNITER FOR A SOLID PROPELLANT ROCKET ENGINE, SAID ENGINE HAVING A COMBUSTION CHAMBER CONFINING A SOLID PROPELLANT, SAID PROPELLANT HAVING A PASSAGE EXTENDING THERETHROUGH CO-AXIALLY WITH THE AXIS OF THE ENGINE, SAID PASSAGE EXTENDING THE LENGTH OF THE ENGINE SO AS TO DEFINE AN INTERNAL BURNING CHARGE, SAID IGNITER COMPRISING: (A) AN OPEN-ENDED CYLINDRICAL CHAMBER POSITIONED AT ONE END OF SAID PASSAGE, SAID OPEN END OF SAID CHAMBER BEING DIRECTED DOWN THE PASSAGE; (B) A MATRIX OF A SOLID ROCKET PROPELLANT COMPRISING A MIXTURE OF A FUEL BINDER AND AN OXIDIZER HAVING DISPERSED THEREIN SOLID PELLETS, SAID PELLETS COMPRISING FINELY DIVIDED METAL AND AN OXIDIZER THEREFOR, ALL OF SAID PELLETS BEING DISPERSED WITHIN AND SURROUNDING BY SAID MATRIX, SAID MATRIX BEING CONFINED BY SAID CYLINDRICAL OPEN-ENDED CHAMBER, SAID MATRIX HAVING A PASSAGE EXTENDING THERETHROUGH CO-AXIALLY WITH THE AXIS OF THE ENGINE; (C) IGNITION MEANS FOR SAID MATRIX POSITIONED AT ONE END THEREOF; AND (D) SAID IGNITOR PROVIDING A SUBSONIC FLOW OF HOT GAS HAVING HOT METALLIC PARTICLES THEREIN FOR THE IGNITION OF THE SOLID PROPELLENT OF THE ROCKET ENGINE.
US54966A 1960-09-09 1960-09-09 Rocket engine igniter Expired - Lifetime US3129561A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
US54966A US3129561A (en) 1960-09-09 1960-09-09 Rocket engine igniter

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US54966A US3129561A (en) 1960-09-09 1960-09-09 Rocket engine igniter

Publications (1)

Publication Number Publication Date
US3129561A true US3129561A (en) 1964-04-21

Family

ID=21994676

Family Applications (1)

Application Number Title Priority Date Filing Date
US54966A Expired - Lifetime US3129561A (en) 1960-09-09 1960-09-09 Rocket engine igniter

Country Status (1)

Country Link
US (1) US3129561A (en)

Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3273335A (en) * 1964-05-04 1966-09-20 Edward S Gravlin Manifold ignition system for solid propellant rockets
US3292545A (en) * 1963-10-01 1966-12-20 Mitsubishi Heavy Ind Ltd Propellant grain
US3613597A (en) * 1964-01-16 1971-10-19 North American Rockwell Solid propellant grain
US3811380A (en) * 1958-01-30 1974-05-21 Cava Ind Rocket and propellant therefor
JPS5011854Y1 (en) * 1970-06-26 1975-04-12
US3878791A (en) * 1969-01-17 1975-04-22 Us Air Force Encapsulated pellet igniter charge
FR2322270A1 (en) * 1975-08-25 1977-03-25 Secr Defence Brit IGNITER FOR ROCKET ENGINE
US4080901A (en) * 1976-04-20 1978-03-28 The United States Of America As Represented By The Aministrator Of The National Aeronautics And Space Administration Molded composite pyrogen igniter for rocket motors
US4573316A (en) * 1983-03-16 1986-03-04 Her Majesty The Queen In Right Of Canada, As Represented By The Minister Of National Defence Of Her Majesty's Canadian Government Integrated weatherseal/igniter for solid rocket motor
US4630539A (en) * 1985-06-24 1986-12-23 The United States Of America As Represented By The Secretary Of The Army Device for flash suppression of a rocket motor
FR2702824A1 (en) * 1993-03-15 1994-09-23 Lacroix E Tous Artifices Igniter unit especially for lighting a gas for cooling a propulsion unit
US20030233956A1 (en) * 2002-05-15 2003-12-25 Atlantic Research Corporation Destroying airborne biological and/or chemical agents with solid propellants

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2906094A (en) * 1954-04-14 1959-09-29 Glenn H Damon Fuel and rapid ignition apparatus for ignition of fuel in ram jets and rockets
US2914910A (en) * 1954-10-11 1959-12-01 United Aircraft Corp Propyl nitrate monofuel and method of use with nickel salts
US2935839A (en) * 1958-02-03 1960-05-10 Ethyl Corp Igniter compositions
US2955535A (en) * 1958-05-28 1960-10-11 Olin Mathieson Ignition assembly for perforated cylindrical charge
US2959001A (en) * 1957-10-09 1960-11-08 American Potash & Chem Corp Ignition of rocket thrust devices
US2980021A (en) * 1956-06-14 1961-04-18 Phillips Petroleum Co Ignition of solid rocket propellants
US2990683A (en) * 1957-12-30 1961-07-04 Phillips Petroleum Co Ignition of solid rocket propellants

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2906094A (en) * 1954-04-14 1959-09-29 Glenn H Damon Fuel and rapid ignition apparatus for ignition of fuel in ram jets and rockets
US2914910A (en) * 1954-10-11 1959-12-01 United Aircraft Corp Propyl nitrate monofuel and method of use with nickel salts
US2980021A (en) * 1956-06-14 1961-04-18 Phillips Petroleum Co Ignition of solid rocket propellants
US2959001A (en) * 1957-10-09 1960-11-08 American Potash & Chem Corp Ignition of rocket thrust devices
US2990683A (en) * 1957-12-30 1961-07-04 Phillips Petroleum Co Ignition of solid rocket propellants
US2935839A (en) * 1958-02-03 1960-05-10 Ethyl Corp Igniter compositions
US2955535A (en) * 1958-05-28 1960-10-11 Olin Mathieson Ignition assembly for perforated cylindrical charge

Cited By (16)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3811380A (en) * 1958-01-30 1974-05-21 Cava Ind Rocket and propellant therefor
US3292545A (en) * 1963-10-01 1966-12-20 Mitsubishi Heavy Ind Ltd Propellant grain
US3613597A (en) * 1964-01-16 1971-10-19 North American Rockwell Solid propellant grain
US3273335A (en) * 1964-05-04 1966-09-20 Edward S Gravlin Manifold ignition system for solid propellant rockets
US3878791A (en) * 1969-01-17 1975-04-22 Us Air Force Encapsulated pellet igniter charge
JPS5011854Y1 (en) * 1970-06-26 1975-04-12
FR2322270A1 (en) * 1975-08-25 1977-03-25 Secr Defence Brit IGNITER FOR ROCKET ENGINE
US4080901A (en) * 1976-04-20 1978-03-28 The United States Of America As Represented By The Aministrator Of The National Aeronautics And Space Administration Molded composite pyrogen igniter for rocket motors
US4573316A (en) * 1983-03-16 1986-03-04 Her Majesty The Queen In Right Of Canada, As Represented By The Minister Of National Defence Of Her Majesty's Canadian Government Integrated weatherseal/igniter for solid rocket motor
US4630539A (en) * 1985-06-24 1986-12-23 The United States Of America As Represented By The Secretary Of The Army Device for flash suppression of a rocket motor
FR2702824A1 (en) * 1993-03-15 1994-09-23 Lacroix E Tous Artifices Igniter unit especially for lighting a gas for cooling a propulsion unit
US20030233956A1 (en) * 2002-05-15 2003-12-25 Atlantic Research Corporation Destroying airborne biological and/or chemical agents with solid propellants
US6748868B2 (en) * 2002-05-15 2004-06-15 Atlantic Research Corp. Destroying airborne biological and/or chemical agents with solid propellants
US20040112487A1 (en) * 2002-05-15 2004-06-17 Miskelly Hermann L. Solid propellant formulations and methods and devices employing the same for the destruction of airborne biological and/or chemical agents
US6782827B2 (en) * 2002-05-15 2004-08-31 Aerojet-General Corporation Solid propellant formulations and methods and devices employing the same for the destruction of airborne biological and/or chemical agents
US6808572B2 (en) * 2002-05-15 2004-10-26 Aerojet-General Corporation Solid propellant formulations and methods and devices employing the same for the destruction of airborne biological and/or chemical agents

Similar Documents

Publication Publication Date Title
US2703960A (en) Rocket
US3062147A (en) Igniter for solid propellant grains
US4892037A (en) Self consumable initiator
US2959001A (en) Ignition of rocket thrust devices
US2561670A (en) Ignitor
US3129561A (en) Rocket engine igniter
US6846372B1 (en) Reactively induced fragmentating explosives
US2478958A (en) Pressure release
US2942547A (en) Gas generating assembly
US2597641A (en) Pressure-operated starting device
US4671179A (en) Cartridged ammunition for gun barrel weapons
US3003419A (en) Rod-type pyrogenic igniter
US3392673A (en) Consumable pyrogen igniter
US3143853A (en) Solid propellant burn area control
US3021786A (en) Blasting device
US3754511A (en) Fuel and fuel igniter for ram jet and rocket
US3132475A (en) Hybrid rocket propulsion system
US4208967A (en) Squib design
US3667396A (en) Solid propellant grain igniter
US2973713A (en) Ignition of solid rocket propellants
Mitsuno et al. Combustion of metallized propellants for ducted rockets
US3000306A (en) Solid propellant propulsion system
US3976008A (en) Pyrotechnic seeding pellet
US4391196A (en) Add-on igniter for pyrogen type igniter
US2917894A (en) Operation of solid propellant rockets